XFOIL         Version 6.94
  
 Calculated polar for: USA 29 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4588   0.01217   0.00442  -0.0938  0.6275  0.3398
   0.500   0.5057   0.01248   0.00465  -0.0918  0.6137  0.3575
   1.000   0.5541   0.01279   0.00488  -0.0903  0.6007  0.3749
   1.500   0.6024   0.01293   0.00499  -0.0888  0.5884  0.3908
   2.000   0.6524   0.01304   0.00509  -0.0877  0.5765  0.4043
   2.500   0.7006   0.01308   0.00515  -0.0863  0.5648  0.4166
   3.000   0.7494   0.01322   0.00530  -0.0850  0.5531  0.4285
   3.500   0.7734   0.01289   0.00493  -0.0781  0.5205  0.4384
   4.000   0.7959   0.01293   0.00482  -0.0712  0.4879  0.4473
   4.500   0.8260   0.01295   0.00491  -0.0659  0.4703  0.4570
   5.000   0.8445   0.01322   0.00511  -0.0584  0.4357  0.4665
   5.500   0.8731   0.01348   0.00544  -0.0532  0.4150  0.4833
   7.000   1.0133   0.02157   0.01260  -0.0600  0.0068  1.0000
   7.500   1.0354   0.02290   0.01403  -0.0552  0.0065  1.0000
   8.000   1.0555   0.02448   0.01574  -0.0505  0.0065  1.0000
   8.500   1.0728   0.02638   0.01779  -0.0459  0.0066  1.0000
   9.000   1.0869   0.02869   0.02026  -0.0414  0.0068  1.0000
   9.500   1.0963   0.03157   0.02333  -0.0370  0.0070  1.0000
  10.000   1.0996   0.03525   0.02722  -0.0329  0.0072  1.0000
  10.500   1.0976   0.03979   0.03195  -0.0294  0.0075  1.0000
  11.000   1.0923   0.04511   0.03745  -0.0266  0.0077  1.0000
  11.500   1.0821   0.05129   0.04379  -0.0242  0.0079  1.0000
  12.000   1.0777   0.05700   0.04963  -0.0223  0.0082  1.0000
  12.500   1.0841   0.06166   0.05446  -0.0205  0.0088  1.0000
  13.000   1.0936   0.06537   0.05824  -0.0172  0.0098  1.0000